Large gas turbine engines are widely used for aircraft propulsion and for ground based power generation. Such large gas turbine engines are of the axial type, and include a compressor section, a combustor section, and a turbine section, with the compressor section normally preceded by a fan section. An annular flow path for working medium gases extends axially through the sections of the engine. Each of the fan, compressor, and turbine sections comprises a plurality of disks mounted on a shaft, with a plurality of airfoil shaped blades projecting radially from the disks. A hollow case surrounds the various engine sections. A plurality of stationary vanes are located between the disks and project inwardly from the case assembly which surrounds the disks.
During operation of the fan, compressor, and turbine sections, as the working medium gases are flowed axially, they alternately contact moving blades and the stationary vanes. In the fan and compressor sections, air is compressed and the compressed air is combined with fuel and burned in the combustion section to provide high pressure, high temperature gases. The working medium gases then flow through the turbine section, where energy is extracted by causing the bladed turbine disks to rotate. A portion of this energy is used to operate the compressor section and the fan section.
Engine efficiency depends to a significant extent upon minimizing leakage of the gas flow to maximize interaction between the gas stream and the moving and stationary airfoils. A major source of inefficiency is leakage of gas around the tips of the compressor blades, between the blade tips and, the engine case. Accordingly, means to improve efficiency by reduction of leakage are increasingly important. Although a close tolerance fit may be obtained by fabricating the blade tips and the engine case to mate to a very close tolerance range, this fabrication process is extremely costly and time consuming. Further, when the assembly formed by mating the blade tips and the engine case is exposed to a high temperature environment and rotational forces, as when in use, the coefficients of expansion of the blade tips and the engine case parts may differ, thus causing the clearance space to either increase or decrease. A significant decrease in clearance results in contact between blades and housing, and friction between the parts generates heat causing a significant elevation of temperatures and possible damage to one or both members. On the other hand, increased clearance space would permit gas to escape between the compressor blade and housing, thus decreasing efficiency.
One approach to increase efficiency is to apply an abradable coating of suitable material to the interior surface of the compressor housing, which when abraded allows for the creation of a channel between the blade tips and the housing. Leakage between the blade tips and the housing is limited to airflow in the channel. Various coating techniques have been employed to coat the inside diameter of the compressor housing with an abradable coating that can be worn away by the frictional contact of the compressor blade, to provide a close fitting channel in which the blade tip may travel. Thus, when subjecting the coated assembly to a high temperature and stress environment, the blade and the case may expand or contract without resulting in significant gas leakage between the blade tip and the housing.
However, it is critical that the blade tips not degrade when contacted with the coatings applied to the interior surface of the compressor housing. To increase the durability of the blade tips which rub against the abradable seals, abrasive layers are sometimes applied to the blade tip surface by a variety of methods. See, for example, U.S. Pat. No. 4,802,828, of Rutz et al, which suggests several techniques for providing an abrasive layer on a blade tip, including powder metallurgy techniques, plasma spray techniques, and electroplating techniques; Schaefer et al, U.S. Pat. No. 4,735,656, which teaches application of an abrasive comprising ceramic particulates in a metal matrix by controlled melting and solidification of the matrix metal; or, Schaefer et al, U.S. Pat. No. 4,851,188, which teaches a sintering operation for application of an abrasive layer to the tip of a superalloy gas turbine blade.
Plasma spraying devices and techniques are well known in the art for depositing protective coatings on underlying substrates. One known device is illustrated in U.S. Pat. No. 3,145,287 to Siebein et al entitled "Plasma Flame Generator and Spray Gun ". In accordance with the teaching of the Siebein et al patent, a plasma-forming gas forms a sheath around an electric arc. The sheath of gases constricts and extends the arc part way down the nozzle. The gas is converted to a plasma state and leaves the arc and nozzle as a hot plasma stream. Powders are injected into the hot plasma stream and propelled onto the surface of the substrate to be coated.
U.S. Pat. Nos. 3,851,140 to Coucher entitled "Plasma Spray Gun and Method for Applying Coatings on a Substrate" and 3,914,573 to Muehlberger entitled "Coating Heat Softened Particles by Projection in a Plasma Stream of Mach 1 to Mach 3 Velocity" disclose contemporaneous coating technology.
This above art notwithstanding, scientists and engineers working under the direction of Applicant's assignee are seeking to improve the process of applying thermal spray coating to substrates in a gas turbine engine. In particular, they have sought to improve the application time of the thermal spray coating using the plasma spraying devices, and to produce a process that is tolerant of variations in flow parameters affecting the spray coating.